The present invention relates to a method of repairing single crystal nickel base superalloy components such that mechanical property degradations are minimized and repair joint strength is maximized.
Superalloys are used for various components in gas turbine engines where the components are exposed to high temperatures and pressures for extended periods of time. A typical application is high pressure turbine vanes, where the temperature of the component can rise to about 2000xc2x0 F. Under these conditions, the component is expected to retain its shape and strength long enough to provide economical operation of the engine without unduly frequent service or replacement requirements.
Typically, superalloy components have been produced by precision casting, which provides a near-net shape component requiring only minimal machining operations to reach final configuration. To provide additional protection for the material from the high temperatures and corrosive environment due to combustion processes, a protective coating is commonly applied. Components with adequate properties for the desired applications have been successfully produced with no heat treatment other than the heating cycle incurred as part of the coating process, which is effectively a precipitation heat treatment of the as-cast material.
After extended service, some of these components incur damage, due to erosion, thermal mechanical fatigue-induced cracking caused by the frequent cycling between ambient and operating temperatures, or creep, which causes the airfoil portions of the components to bow or the platform portions of the component to twist away from their original positions, with a resultant change in the operational characteristics of the components. A number of methods for repairing these components have been developed and used.
One such repair method is illustrated in U.S. Pat. Nos. 5,549,767 and 5,922,150, both to Pietruska et al. This method for repairing cobalt-based superalloy gas turbine engine components comprises applying a mixture of a base alloy powder and a base alloy powder with a melting point depressant to the surface of the component and heating at a temperature in the range of 2250xc2x0 F. to 2300xc2x0 F. to diffuse the melting point depressant isothermally into the base alloy. A protective coating is then applied, during which a heating cycle which ages the base material is used.
U.S. Pat. No. 5,741,378 to Pietruska et al. illustrates a method for restoring the mechanical properties of carbide-containing cobalt-based superalloy gas turbine engine components. The method includes solution heat treating to a temperature in the range of 2250xc2x0 F. to 2300xc2x0 F. for one to twelve hours to dissolve complex carbides and aging at approximately 1965xc2x0 F. to 1975xc2x0 F. for two to twenty four hours.
Braze-type repairs to non-single crystal nickel base superalloy components used in gas turbine engines are typically accomplished using a heat treatment method that re-solutions the alloy to assure no mechanical property debits. At the same time, this heat treatment assures maximum diffusion of the repair zone without any incipient melting in the repair zone or the base alloy. This method is effective for non-single crystal nickel based superalloys; however, for single crystal nickel based superalloys, this method results in significant recrystallization that greatly reduces fatigue properties and also results in an overdiffused repair zone with local areas of melting.
Thus, there remains a need for an effective method for repairing components formed from single crystal nickel based superalloys.
Accordingly, it is an object of the present invention to provide an improved method for repairing single crystal nickel based superalloy components so as to minimize mechanical property degradations.
It is a further object of the present invention to provide a method as above which maximizes joint strength.
The foregoing objects are attained by the method of the present invention.
In accordance with the present invention, a method for repairing components formed from a single crystal nickel based superalloy broadly comprises applying a repair alloy to a single crystal nickel based superalloy component to be repaired and heating said component and said applied repair alloy to a temperature that avoids recrystallization and repair zone incipient melting of the single crystal nickel based superalloy. In a first embodiment of the method of the present invention, the heating step comprises heat treating the component and the applied repair alloy at a temperature in the range of from about 2150xc2x0 F. to about 2275xc2x0 F., preferably about 2200xc2x0 F., for a time period in the range of from about 5 hours to 24 hours, preferably about 10 hours. Following heating, the component is rapidly cooled and subjected to an aging treatment.
In a second embodiment of the present invention, the heating step comprises heating the component and the applied repair alloy to a temperature in the range of from about 2250xc2x0 F. to about 2350xc2x0 F., preferably about 2300xc2x0 F., holding the component and the applied repair alloy at this temperature for a time in the range of less than about 30 minutes, preferably 15 minutes or less, rapidly cooling the component and the applied repair alloy to a temperature in the range of from about 1950xc2x0 F. to about 2075xc2x0 F., preferably about 2025xc2x0 F., and holding the component and the applied repair alloy at this temperature for a time in the range of from about 5 hours to about 24 hours, preferably about 10 hours. Following this heat treatment, the component is rapidly cooled and subjected to an aging treatment.
Other details of the repair method of the present invention, as well as other objects and advantages attendant thereto, are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements.